The present invention relates to gas turbine engines, and more particularly, to an improved turbine shroud assembly.
Gas turbine aircraft engines comprising a compressor portion, a combustion chamber and an axial turbine portion are well known. The heat generated during combustion, however, nevertheless presents challenges when designing structural elements that are to be exposed to the high combustion temperatures, such as the elements of the turbine portion of the engine, where temperatures can easily reach 2000xc2x0 F. The task of designing components capable of withstanding such elevated temperatures is made additionally difficult by the need to keep weight to a minimum. Weight reduction of gas turbine engines used for aircraft applications is becoming increasingly important, and as such weight considerations remain a critical design focus for new gas turbine engine components.
Two main approaches have been taken to meet these requirements for turbine components. The first comprises using various fluid cooling systems, often using air as the cooling fluid, to reduce the peak temperatures of the metal turbine structure, without requiring a reduction in turbine inlet temperatures which would negatively affect overall engine performance. The problem with this approach is that the cooling air is extracted from air that could otherwise be used for the propulsion cycle, thereby reducing the engine performance. The higher the turbine inlet temperature, the more cooling air is required to maintain the turbine components at acceptable temperatures, and therefore the more air is required to be extracted from the working air.
The second approach taken to deal with high turbine operating temperatures is to use components made of materials capable of withstanding higher temperatures with little or no fluid cooling required. As such, ceramics have become more and more utilized for their ability to effectively withstand high temperatures without negative affects on its material strength. Ceramic as a material is additionally attractive for use in aircraft applications, because of its relatively low weight in comparison to traditionally used metals and metal alloys.
However, certain characteristics of ceramic materials prevent direct replacement of metal alloy turbine components with ceramic components. Ceramic materials are generally much more brittle and have lower tensile strength than most metals. A major obstacle restricting the use of ceramic components in high temperature regions of gas turbine engines is the considerable difference in thermal expansion of ceramic materials in comparison to metals or metal alloys. The thermal expansion coefficients of ceramic materials are only a small fraction of those of conventionally used nickel alloy materials, for example. This presents considerable difficulties when a ceramic element and a metal alloy element are interfaced.
Some attempts have been made to solve this thermal growth mismatch problem when using ceramic turbine components in gas turbine engines. U.S. Pat. No. 4,087,199, issued May 2, 1978 to Hemsworth et al., for example, discloses a ceramic turbine shroud assembly comprising a plurality of ceramic blocks which are arranged in a ring around the tips of the rotating turbine blades. Each ceramic block is provided with a pair of dovetail surfaces formed on opposite sides of the block which function as wedging surfaces. Metallic clamping means in the form of a pair of annular spring-like members, hold the blocks in the assembly and produce a preloaded radial force against the dovetail surfaces. This preloaded clamping of the blocks against the rigid stops establishes the shroud in the proper radial position, but does not permit the shroud to be resiliently, eccentrically displaced.
U.S. Pat. No. 3,146,992, issued to Farrell Sep. 1, 1964, also discloses a turbine shroud support structure. Farrell does not teach the use of a ceramic shroud, but provides a sprung shroud designed to maintain clearances between the turbine blade tips and the shroud. The turbine shroud support structure comprises bimetallic thermal support strips which are provided for maintaining the desired clearances between a circumferentially extending segmented shroud ring and the tips of a row of turbine blades. The bimetallic support strips are supported by their ends in the space between the segmented shroud ring and casing, each strip positioned with its layer having the lowest coefficient of expansion adjacent the casing. The unsupported center of each bimetallic support strip is connected to a respective shroud ring segment. With increasing operating temperature, the bimetallic strips deflect to move the shroud inwardly relative to the exterior casing. With decreasing operating temperatures, they deflect to move the shroud outwardly relative to the casing.
Both of these references, however, disclose segmented turbine shrouds. Segmented shrouds are less efficient for sealing purposes in comparison with continuous shroud rings, and permit more hot gas leakage between the shroud segments. Additionally, segmented rings create greater difficulty in setting turbine blade tip clearances, exact shroud diameter and roundness.
It is an object of the present invention to provide an improved turbine shroud assembly.
It is another object of the present invention to provide a ceramic turbine shroud ring and a mounting method thereof.
It is yet another object of the present invention to provide an attachment for a ceramic turbine shroud in a metal housing such that the thermal expansion difference between the shroud ring and the support housing is compensated.
Therefore, in accordance with one the present invention, there is provided a shroud assembly for a turbine portion of a gas turbine engine, the shroud assembly comprising: an annular ceramic shroud ring, circumferentially disposed about radially extending blades of a turbine rotor and partially defining an annular hot gas passage of said turbine portion; a plurality of arcuate shroud support segments, radially disposed outwardly of said ceramic shroud ring and contiguous therewith; a plurality of inwardly biased resilient members, each engaged between one of said shroud support segments and an outer annular turbine support case composed of a material having a different thermal expansion coefficient than said ceramic shroud ring, said resilient members maintaining contact between said shroud support segments and said ceramic shroud ring; and said shroud supporting segments and said resilient members being adapted to deflect to compensate for relative thermal growth differences between said ceramic shroud ring and said turbine support case.
In accordance with a second aspect of the present invention, a ceramic shroud assembly is provided for a gas turbine engine turbine portion comprising a turbine rotor having radially extending turbine blades, the ceramic shroud assembly comprising: a continuously uninterrupted ceramic shroud ring, circumferentially disposed about said turbine blades and partially defining an annular hot gas passage of said turbine portion; whereby said continuously uninterrupted ceramic shroud ring minimizes hot gas leakage from tips of said turbine blades.
In accordance with a third aspect of the present invention, a shroud assembly is provided for a turbine portion of a gas turbine engine, the shroud assembly comprising: a turbine shroud, circumferentially disposed about radially extending blades of a turbine rotor and partially defining an annular hot gas passage of said turbine portion; a plurality of arcuate shroud support segments, radially disposed outwardly of said turbine shroud and contiguous therewith; a plurality of inwardly biased leaf springs, each engaged between one of said shroud support segments and an outer annular turbine support case, said leaf springs maintaining contact between said shroud support segments and said turbine shroud; said turbine shroud and said outer annular turbine support case being composed of materials having different thermal expansion coefficients; whereby said shroud supporting segments and said leaf springs being adapted to deflect to compensate for relative thermal growth differences between said turbine shroud and said turbine support case.